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Marco Polo: a Sample Return Mission to a Primitive NEO

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Marco Polo: a Sample Return Mission to a Primitive NEO Marco Polo: A sample return mission to a primitive NEO Session IX: Sample Return Challenges David Agnolon ESA-ESTEC, Directorate of Science & Robotic Exploration Email: [email protected] P. Coste, R. Drai, D. Escorial, P. Falkner, L. Ferracina, D. Koschny, H. Ritter, J. Romstedt (ESA-ESTEC) M. Khan (ESA-ESOC) A. Barucci (Observatory Of Paris) 6th International Planetary Probe Workshop ; 21-06-2008 1 Outline Cosmic-Vision context Marco Polo science requirements Scope of the CDF study NEO target selection and baseline mission profile Sampling strategy, sample acquisition and transfer system Orbiting, descent and landing on a NEO High-speed Earth re-entry Overall spacecraft design Technology development approach 6th International Planetary Probe Workshop ; 21-06-2008 2 ESA Cosmic-Vision Future of the ESA Scientific programme 2017-202X Cosmic Vision Call for Proposals in April 2007 for: • 1 “medium M” class mission for launch in 2017 • 1 “large L” class mission for launch in 2018 50 proposals receivedNEO Sample return mission Æ 9 proposals selected (i.e. 7 missions) by ESA advisory structures assessment phaseEuropean for 1st cycle Japanese of Cosmic-Vision team (950 M€ budget) Cross-Scale Lead proposers: Spica • A. Barucci (LESIA, Paris Euclid Laplace Observatory) Marco Polo Tandem • M. Yoshikawa (JSPEC/JAXA) Plato XEUS Down selection of 2 M missions to enter definition phase ~ mid-2009 Final M mission selected in 2011 for implementation 6th International Planetary Probe Workshop ; 21-06-2008 3 Requirements and payload Defined by US, European and Japanese scientists Go to a D, T, C or B type NEO (most primitive) Return > 30 g sample (goal 100 g) Place sample in their global/local context Provide complementary info not available from the samples Multiple sampling locations (non-uniform composition) Maximum sample temperature: + 40oC (organics) Sample composition (cm-sized pebbles + small particles) Avoid contamination of the Spatial resolution Spatial Spatial sample for imaging in the resolution for resolution for visual VIS/IR mid-IR Instruments: wide & narrow spectrometer instrument angle camera, Vis-NIR & Global Order of dm Order of m Order of mid-IR spectrometer, laser characterisation 10 m altimeter, neutral particle Local x 5 Order of mm Order of dm Order of dm characterisation analyzer, radio science Context Tens of μm - - experiment measurements 6th International Planetary Probe Workshop ; 21-06-2008 4 Scope of the ESA CDF Study Internal ESA pre-assessment study (similar to NASA JPL Team X) Objectives: • Critical assessment of initial Marco Polo proposal • Establish a feasible & cost-efficient mission profile option • Fulfil science objectives & requirements • Feed forward upcoming industry studies • Help define required technology development Discussions with JAXA ongoing Æ As a starting point CDF study focuses on European expertise : NOT meant to be representative of selected Phase A scenario M-mission Cosmic-Vision technology constraints: • No major development can be undertaken until definition phase unless strong generic interest • TRL 5/6 by 2011 Æ build on ongoing development and off-the-shelf equipment 6th International Planetary Probe Workshop ; 21-06-2008 5 NEO selection D (km) Provisional q assumes Preliminary global screening designation Number Type (AU) Q (AU) i (deg) HV p=0.06 Prot (hrs) 1989 UQ 65679 C 0.67 1.16 1.3 19.28 0.76 7.733 (> 5000 known NEO!) 1999 JU3 162173 Cg 0.96 1.42 5.9 19.2 0.78 7.5 2001 SK162 162998 T 1.01 2.84 1.7 17.77 1.52 68 Global optimization algorithms 2001 SG286 D 0.89 1.83 7.8 20.93 0.35 ? 1999 RQ36 101955 B or C1 0.9 1.36 6 20.81 0.37 2.146 run over ~ 15 targets based on 1996 FG3 C 0.69 1.42 2 18.22 1.23 3.5942 2001 AE2 NEW138911 T 1.24TARGET 1.46 1.66 19 0.86 ? specific optimization criteria 2001 FC7 68278 Cg1 1.27 1.6 2.6 18.17 1.26 ? 1977 VA 136564 C or Xc1 1.13 2.6 3 19.06 0.84 ? More in-depth mission analysis + 1998 KU2 152679PROPOSALS Cb 1.01 3.5 4.9 16.47 2.76 ? 1960 UA = design iterations for 4 targets Anza 2061 TCG 1.05 3.48 3.8 16.56 2.65 5.75, 11.5 1979 VA = Wilson- WELCOMEDormant !!! Harrington 4015 comet 0.99 4.28 2.8 15.99 3.44 3.56, 6.1 1991 DB 14402 C 1.03 2.41 11.4 18.4 1.13 2.266 2000 RW37 162567 C 0.94 1.56 13.7 19.74 0.61 ? V1 +Δ 2× Δ 2 + ΔVV i 1998 UT18 85774 C 0.94 1.87 13.6 19.07 0.83 34 1 To be confirmed by observations ΔV2 ΔV1 NEO LAUNCH CRUISE ARRIVAL SCI DEPARTURE CRUISE ENTRY TOTAL name escape V-inf dec swb1 swb2 swb3 date kg yea date swb date V-ent kg m/s years 19/11/30 4.000 -10 20/12/04 22/01/20 1550 1.5 23/07/18 24/12/03 12.13 1303 2665 5.0 1999 JU3 17/12/28 3.466 10 20/08/17 20/12/05 22/04/28 1759 1.2 23/07/18 24/12/04 12.13 1471 2288 6.9 18/08/27 3.394 -45 19/02/13 20/12/10 23/03/07 1294 0.8 23/12/06 26/01/08 12.75 1188 2951 7.4 2001 SK162 17/08/25 3.396 0 18/08/25 19/02/12 20/12/09 23/04/04 1444 0.7 23/12/06 26/01/08 12.75 1324 2615 8.4 18/09/19 3.490 -30 19/03/05 20/12/07 1610 1.7 22/08/04 23/06/24 23/11/16 11.83 1366 2517 5.2 1989 UQ 17/09/20 3.492 0 18/09/20 19/03/05 20/12/20 1702 1.6 22/08/04 23/06/24 23/11/16 11.83 1444 2345 6.2 18/01/28 3.889 6 19/03/27 20/12/25 22/07/19 1041 1.3 23/10/24 25/05/22 15.13 903 3803 7.3 2001 SG286 18/03/24 3.553 0 19/04/14 20/05/21 22/03/22 1380 1.6 23/10/24 25/05/22 15.13 1195 2932 7.2 Venus Earth Mars 6th International Planetary Probe Workshop ; 21-06-2008 6 Baseline mission analysis Launch by Soyuz-Fregat 2-1b from Kourou on direct escape (V = 3.49 km.s-1, Dec=0o) inf 10 months waiting September March 1566 kg launch mass 2019 period Æ 2018 October capability 2021 Total mission Arrival: December -1 ΔV ~ 1143 m.s 2020 Oct. 2021 Æ Re-entry velocity July 2022 = -1 September 9 months ~ 11.8 km.s 2017 science and sampling Mission duration 6.2 November operations years, incl. 1.6 years at 2023 NEO June 2023 Departure: July 2022 6th International Planetary Probe Workshop ; 21-06-2008 7 Proximity operations Both uncontrolled (for radio science) and controlled orbits have been analyzed assuming: • 1989 UQ physical properties (tumbling, 4:2:1 shape, ~ 760 m diameter, 7.7 h rotation, 1300 kg/m 3, etc. ) • Solar radiation pressure • Sun’s gravity influence (influence of planets’ negligible) 1 month 7 months 6th International Planetary Probe Workshop ; 21-06-2008 8 Sampling strategy, sample acquisition and transfer system Possible strategies: • Hover & go • Touch & go • Short-term landing • Long-term landing 1m Rotating corer used as Image courtesy: JAXA sample vessel Mounted at the tip of telescopic boom Inserted into rotating corer holder (ERC backcover) further transferred Elevator to ERC via elevator Corer holder 20 N down-thrust during sampling operations Corer 20-30 minutes operations Telescopic Arm Æ Sampling Debris Smooth landing hatch Cover 6th International Planetary Probe Workshop ; 21-06-2008 9 Descent and landing Navigation strategy (vision-based) • Get a good shape/gravity model and map hazards from orbit • Descent down to 150 m • “Go” decision Æ Autonomous + 90o slew • Controlled descent (lateral) -1 • Landing conditions: VV < 20 cm.s , VH < 5 cm.s -1, θ < 10o • Landing accuracy < 5 m Æ safe site • Autonomous battery-powered descent, landing, sampling and ascent operations ( ~ Image courtesy: Astrium Ltd 2 hours) Landing structure/mechanism • Philae damping system or multistage crushable • Much higher clearance requirement • No anchoring 6th International Planetary Probe Workshop ; 21-06-2008 10 High-speed Earth re-entry 11.8 km.s-1, FPA~-12o Landing load on the sample: 800 g Peak heat flux: 11.3 MW.m-2 Heat load: 209 MJ.m-2 Entry duration: 484 s Sample container Rear TPS (Norcoat Liege) -3 Outer sphere PICA-like (260 kg.m ) front (FRP-Kevlar) Sample containment structure TPS, Norcoat-Liege rear TPS (FRP-Kevlar) 45o half-cone angle Shape foam 1.1 m base diameter Main structure (CFRP) 200 mm RVC foam Damping foam (RVC) 76 kg capsule Front TPS (PICA type carbon phenolic ablator) 6th International Planetary Probe Workshop ; 21-06-2008 11 Spacecraft design Outbound Descent Sampling Earth Return Earth Re- propulsion Orbiter Module Module Vehicle entry Capsule Module Body-mounted solar arrays (485 W) Low-thrust dual mode propulsion system 3-axis controlled 3 landing legs Central SATS accommodation Surface thermal control to be further analyzed Mass (i.e. keep sample Orbiter-lander dry mass incl. system margin 646 kg cold & high Orbiter-lander wet mass 1191 kg ERC 76 kg temperature Orbiter-lander propellant mass 545 kg mechanisms) Launch mass 1267 kg Launch mass incl.
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